Gas turbine engines, such as those used to power modern aircraft or in industrial applications, include a compressor for pressurizing a supply of air, a combustor for burning a hydrocarbon fuel in the presence of the pressurized air, and a turbine for extracting energy from the resultant combustion gases. Generally, the compressor, combustor and turbine are disposed about a central engine axis with the compressor disposed axially upstream of the combustor and the turbine disposed axially downstream of the combustor.
An exemplary combustor features an annular combustion chamber defined between a radially inboard liner and a radially outboard liner extending aft from a forward bulkhead. The radially outboard liner extends circumferentially about and is radially spaced from the inboard liner, with the combustion chamber extending fore to aft therebetween. Exemplary liners are double-wall structured, having an inner heat shield and an outer support shell. Combustion air admission holes penetrate the outboard liner and the inboard liner to admit combustion air into the combustion chamber along the length of the combustion chamber. A plurality of circumferentially distributed fuel injectors and associated swirlers or air passages is mounted in the forward bulkhead. The fuel injectors project into the forward end of the combustion chamber to supply the fuel. The swirlers impart a swirl to inlet air entering the forward end of the combustion chamber at the bulkhead to provide rapid mixing of the fuel and inlet air. Commonly assigned U.S. Pat. Nos. 6,240,731; 6,606,861; 6,810,673 and 7,093,441, the entire disclosures of which are hereby incorporated herein by reference as if set forth herein, disclose exemplary prior art annular combustors for gas turbine engines.
Combustion of the hydrocarbon fuel in air inevitably produces oxides of nitrogen (NOx). NOx emissions are the subject of increasingly stringent controls by regulatory authorities. One combustion strategy for minimizing NOx emissions from gas turbine engines is referred to as rich burn, quick quench, lean burn (RQL) combustion. The RQL combustion strategy recognizes that the conditions for NOx formation are most favorable at elevated combustion flame temperatures, i.e. when the fuel-air ratio is at or near stoichiometric. The combustion process in a combustor configured for RQL combustion has two governing states of combustion: a first state in the forward portion of the combustor that is stoichiometrically fuel-rich and a second state in a downstream portion of the combustor that is stoichiometrically fuel-lean. In conventional practice, all of the fuel is admitted through the plurality of fuel and air admission assemblies disposed in the bulkhead thereby establishing a forward combustion zone in the combustion chamber. A portion of the combustion air is admitted with the fuel into the forward combustion zone, another portion of combustion air is admitted downstream of the forward combustion zone to rapidly quench and dilute the combustion products, and additional air is admitted to the combustion chamber downstream of the quench zone to further dilute the combustion products.
Conventional annular combustors for aircraft gas turbine engines employing the RQL combustion process typically are more effective in reducing NOx emissions during high power engine operating conditions, such as experienced during take-off and climb, than during low and moderate power engine operating conditions, such as experienced during approach, idle, and cruise. At high power, the forward combustion zone is operated fuel-rich, but at low and moderate power, the forward combustion zone is operated closer to stoichiometric fuel to air ratios. Accordingly, a desire exists for a gas turbine combustor that exhibits effective NOx emission reduction performance over the entire range of engine operating conditions, including cruise.